Intercooled cooling air with dual pass heat exchanger

ABSTRACT

A gas turbine engine comprises a main compressor section having a downstream most end, and more upstream locations. A turbine section has a high pressure turbine. A tap taps air from at least one of the more upstream locations in the compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses ng air downstream of the heat exchanger, and delivers air into the high pressure turbine. The heat exchanger has at least two passes, with one of the passes passing air radially outwardly, and a second of the passes returning the air radially inwardly to the compressor. An intercooling system for a gas turbine engine is also disclosed.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a divisional of U.S. patent application Ser. No.14/695,504, filed Apr. 24, 2015.

BACKGROUND

This application relates to improvements in providing cooling air from acompressor section to a turbine section in a gas turbine engine.

Gas turbine engines are known and typically include a fan delivering airinto a bypass duct as propulsion air. Further, the fan delivers air intoa compressor section where it is compressed. The compressed air passesinto a combustion section where it is mixed with fuel and ignited.Products of this combustion pass downstream over turbine rotors drivingthem to rotate.

It is known to provide cooling air from the compressor to the turbinesection to lower the operating temperatures in the turbine section andimprove overall engine operation. Typically, air from the downstreammost end of the compressor has been tapped, passed through a heatexchanger, which may sit in the bypass duct and then delivered into theturbine section. The air from the downstream most end of the compressorsection is at elevated temperatures.

SUMMARY

In a featured embodiment, an intercooling system for a gas turbineengine comprises a heat exchanger for cooling air drawn from a portionof a main compressor section at a first temperature and pressure forcooling the air to a second temperature cooler than the firsttemperature. A cooling compressor compresses air communicated from theheat exchanger to a second pressure greater than the first pressure andcommunicating the cooling air to a portion of a turbine section. Theheat exchanger has at least two passes, with one of the passes passingair radially outwardly, and a second of the passes returning the airradially inwardly to the compressor.

In another embodiment according to the previous embodiment, an auxiliaryfan is positioned upstream of the heat exchanger.

In another embodiment according to the any of the previous embodiments,a main fan delivers bypass air into a bypass duct and into the maincompressor section and the heat exchanger positioned within the bypassduct to be cooled by bypass air.

In another embodiment according to the any of the previous embodiments,the cooling compressor includes a centrifugal compressor impeller.

In another embodiment according to the any of the previous embodiments,the first pass is positioned upstream of the second pass in the bypassduct.

In another embodiment according to the any of the previous embodiments,the first pass is positioned upstream of the second pass in the bypassduct.

In another embodiment according to the any of the previous embodiments,the at least one of the more upstream locations is in a high pressurecompressor.

In another embodiment according to the any of the previous embodiments,the at least one of the more upstream locations is in a low pressurecompressor.

In another embodiment according to the any of the previous embodiments,the at least one of the more upstream locations is in a high pressurecompressor.

In another embodiment according to the any of the previous embodiments,the at least one of the more upstream locations is in a low pressurecompressor.

These and other features may best be understood from the followingspecification and drawings, the following of which is a briefdescription.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2 shows a prior art engine.

FIG. 3 shows one example engine.

FIG. 4 shows a first embodiment heat exchanger.

FIG. 5 shows a second embodiment heat exchanger.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example gas turbine engine 20 thatincludes a fan section 22, a compressor section 24, a combustor section26 and a turbine section 28. Alternative engines might include anaugmenter section (not shown) among other systems or features. The fansection 22 drives air along a bypass flow path B while the compressorsection 24 draws air in along a core flow path C where air is compressedand communicated to a combustor section 26. In the combustor section 26,air is mixed with fuel and ignited to generate a high pressure exhaustgas stream that expands through the turbine section 28 where energy isextracted and utilized to drive the fan section 22 and the compressorsection 24.

Although the disclosed non-limiting embodiment depicts a turbofan gasturbine engine, it should be understood that the concepts describedherein are not limited to use with turbofans as the teachings may beapplied to other types of turbine engines; for example a turbine engineincluding a three-spool architecture in which three spoolsconcentrically rotate about a common axis and where a low spool enablesa low pressure turbine to drive a fan via a gearbox, an intermediatespool that enables an intermediate pressure turbine to drive a firstcompressor of the compressor section, and a high spool that enables ahigh pressure turbine to drive a high pressure compressor of thecompressor section.

The example engine 20 generally includes a low speed spool 30 and a highspeed spool 32 mounted for rotation about an engine central longitudinalaxis A relative to an engine static structure 36 via several bearingsystems 38. It should be understood that various bearing systems 38 atvarious locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatconnects a fan 42 and a low pressure (or first) compressor section 44 toa low pressure (or first) turbine section 46. The inner shaft 40 drivesthe fan 42 through a speed change device, such as a geared architecture48, to drive the fan 42 at a lower speed than the low speed spool 30.The high-speed spool 32 includes an outer shaft 50 that interconnects ahigh pressure (or second) compressor section 52 and a high pressure (orsecond) turbine section 54. The inner shaft 40 and the outer shaft 50are concentric and rotate via the bearing systems 38 about the enginecentral longitudinal axis A.

A combustor 56 is arranged between the high pressure compressor 52 andthe high pressure turbine 54. In one example, the high pressure turbine54 includes at least two stages to provide a double stage high pressureturbine 54. In another example, the high pressure turbine 54 includesonly a single stage. As used herein, a “high pressure” compressor orturbine experiences a higher pressure than a corresponding “lowpressure” compressor or turbine.

The example low pressure turbine 46 has a pressure ratio that is greaterthan about 5. The pressure ratio of the example low pressure turbine 46is measured prior to an inlet of the low pressure turbine 46 as relatedto the pressure measured at the outlet of the low pressure turbine 46prior to an exhaust nozzle.

A mid-turbine frame 58 of the engine static structure 36 is arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 58 further supports bearing systems 38in the turbine section 28 as well as setting airflow entering the lowpressure turbine 46.

Airflow through the core airflow path C is compressed by the lowpressure compressor 44 then by the high pressure compressor 52 mixedwith fuel and ignited in the combustor 56 to produce high speed exhaustgases that are then expanded through the high pressure turbine 54 andlow pressure turbine 46. The mid-turbine frame 58 includes vanes 60,which are in the core airflow path and function as an inlet guide vanefor the low pressure turbine 46. Utilizing the vane 60 of themid-turbine frame 58 as the inlet guide vane for low pressure turbine 46decreases the length of the low pressure turbine 46 without increasingthe axial length of the mid-turbine frame 58. Reducing or eliminatingthe number of vanes in the low pressure turbine 46 shortens the axiallength of the turbine section 28. Thus, the compactness of the gasturbine engine 20 is increased and a higher power density may beachieved.

The disclosed gas turbine engine 20 in one example is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20includes a bypass ratio greater than about six (6), with an exampleembodiment being greater than about ten (10). The example gearedarchitecture 48 is an epicyclical gear train, such as a planetary gearsystem, star gear system or other known gear system, with a gearreduction ratio of greater than about 2.3.

In one disclosed embodiment, the gas turbine engine 20 includes a bypassratio greater than about ten (10:1) and the fan diameter issignificantly larger than an outer diameter of the low pressurecompressor 44. It should be understood, however, that the aboveparameters are only exemplary of one embodiment of a gas turbine engineincluding a geared architecture and that the present disclosure isapplicable to other gas turbine engines.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of pound-mass (lbm) of fuel per hour being burned divided bypound-force (lbf) of thrust the engine produces at that minimum point.

“Low fan pressure ratio” is the pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.50. In another non-limiting embodimentthe low fan pressure ratio is less than about 1.45.

“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram °R)/(518.7° R)]^(0.5). The “Low corrected fan tip speed”, as disclosedherein according to one non-limiting embodiment, is less than about 1150ft/second.

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about 26 fan blades. In anothernon-limiting embodiment, the fan section 22 includes less than about 20fan blades. Moreover, in one disclosed embodiment the low pressureturbine 46 includes no more than about 6 turbine rotors schematicallyindicated at 34. In another non-limiting example embodiment the lowpressure turbine 46 includes about 3 turbine rotors. A ratio between thenumber of fan blades 42 and the number of low pressure turbine rotors isbetween about 3.3 and about 8.6. The example low pressure turbine 46provides the driving power to rotate the fan section 22 and thereforethe relationship between the number of turbine rotors 34 in the lowpressure turbine 46 and the number of blades 42 in the fan section 22disclose an example gas turbine engine 20 with increased power transferefficiency.

Gas turbine engines designs are seeking to increase overall efficiencyby generating higher overall pressure ratios. By achieving higheroverall pressure ratios, increased levels of performance and efficiencymay be achieved. However, challenges are raised in that the parts andcomponents associated with a high pressure turbine require additionalcooling air as the overall pressure ratio increases.

The example engine 20 utilizes air bleed 80 from an upstream portion ofthe compressor section 24 for use in cooling portions of the turbinesection 28. The air bleed is from a location upstream of the downstreamend 82 of the compressor section 24. The bleed air passes through a heatexchanger 84 to further cool the cooling air provided to the turbinesection 54. The air passing through heat exchanger 84 is cooled by thebypass air B. That is, heat exchanger 84 is positioned in the path ofbypass air B. As better described below, a booster compressor 114 helpsdrive the air.

A prior art approach to providing cooling air is illustrated in FIG. 2.An engine 90 incorporates a high pressure compressor 92 downstream ofthe low pressure compressor 94. As known, a fan 96 delivers air into abypass duct 98 and into the low pressure compressor 94. A downstreammost point 82 in the high pressure compressor 92 provides bleed air intoa heat exchanger 93. The heat exchanger is in the path of the bypass airin bypass duct 98, and is cooled. This high pressure high temperatureair from location 82 is delivered into a high pressure turbine 102.

The downstream most point 82 of the high pressure compressor 82 is knownas station 3. The temperature T3 and pressure P3 are both very high.

In future engines, T3 levels are expected to approach greater than orequal to 1350° F. Current heat exchanger technology is becoming alimiting factor as they are made of materials, manufacturing, and designcapability which have difficulty receiving such high temperature levels.

FIG. 3 shows an engine 100 coming within the scope of this disclosure. Afan 104 may deliver air B into a bypass duct 105 and into a low pressurecompressor 106. High pressure compressor 108 is positioned downstream ofthe low pressure compressor 106. A bleed 110 taps air from a locationupstream of the downstream most end 82 of the high pressure compressor108. This air is at temperatures and pressures which are much lower thanT3/P3. The air tapped at 110 passes through a heat exchanger 112 whichsits in the bypass duct 105 receiving air B. Further, the air from theheat exchanger 112 passes through a compressor 114, and then into aconduit 115 leading to a high turbine 117. This structure is all shownschematically.

Since the air tapped at point 110 is at much lower pressures andtemperatures than the FIG. 2 prior art, currently available heatexchanger materials and technology may be utilized. This air is thencompressed by compressor 114 to a higher pressure level such that itwill be able to flow into the high pressure turbine 117.

An auxiliary fan 116 is illustrated upstream of the heat exchanger 112.The main fan 104 may not provide sufficient pressure to drive sufficientair across the heat exchanger 112. The auxiliary fan 116 will ensurethere is adequate air flow in the circumferential location of the heatexchanger 112.

In one embodiment, the auxiliary fan 116 may be variable speed, with thespeed of the fan varied to control the temperature of the air downstreamof the heat exchanger 112. As an example, the speed of the auxiliary fan116 may be varied based upon the operating power of the overall engine.

Further details of the basic system may be found in co-pending patentapplication Ser. No. 14/695,578, entitled “Intercooled Cooling Air,” andfiled on Apr. 24, 2015, which application is hereby incorporated in itsentirety by reference.

Details with regard to an optional arrangement may be found inco-pending patent application Ser. No. 14/695,534, entitled “IntercooledCooling Air With Plural Tap Locations,” and filed on Apr. 24, 2015,which application is hereby incorporated in its entirety by reference.

FIG. 4 shows a unique heat exchanger structure. A bypass duct 120receives the heat exchanger 122. A drive 124 drives compressor impeller125. Impeller 125 is a centrifugal impeller. An inlet 126 receives air,which may be from the low pressure compressor, and delivers that airinto a first heat exchanger pass 128 where it is cooled by the bypassair. The air then turns through an elbow 130 into a second pass 132before reaching the impeller 125. From the impeller 125, the air isdischarged at outlet 134, and then passes to the turbine as in the aboveembodiments.

FIG. 5 shows another embodiment wherein the cooling compressor 140receives relatively hot air from an intermediate location 142 in thehigh pressure compressor 143. The air passes through a first pass 144,an elbow 146, and then a second pass 148 before reaching the compressor140. The air is then delivered into a conduit 150 and passes to the highpressure turbine.

With this second embodiment, the cooling will be more effective,although the air conduits may be somewhat longer.

In both embodiments, the first pass is radially outward and the secondpass is radially inward.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the scope and content of thisdisclosure.

1.-20. (canceled)
 21. An intercooling system for a gas turbine enginecomprising: a heat exchanger for cooling air drawn from a portion of amain compressor section at a first temperature and pressure for coolingthe air to a second temperature cooler than the first temperature; acooling compressor compressing air communicated from the heat exchangerto a second pressure greater than the first pressure and communicatingthe cooling air to a portion of at least a turbine section; and saidheat exchanger having at least two passes, with a first of said passespassing air in a direction having at least a radially outward component,and a second of said passes returning the air in a direction having atleast a radially inward component to the compressor.
 22. Theintercooling system as set forth in claim 21, wherein said first pass ispositioned downstream of said second pass in said bypass duct.
 23. Theintercooling system as set forth in claim 22, wherein said coolingcompressor includes a centrifugal compressor impeller.
 24. Theintercooling system as set forth in claim 22, wherein a main fandelivers bypass air into a bypass duct and into said main compressorsection and said heat exchanger positioned within said bypass duct to becooled by bypass air.
 25. The intercooling system as set forth in claim24, wherein said cooling compressor includes a centrifugal compressorimpeller.
 26. The intercooling system as set forth in claim 24, whereinsaid at least one of said more upstream locations is in a low pressurecompressor.
 27. The intercooling system as set forth in claim 22,wherein said at least one of said more upstream locations is in a lowpressure compressor.
 28. The intercooling system as set forth in claim21, wherein an auxiliary fan is positioned upstream of the heatexchanger.
 29. The intercooling system as set forth in claim 21, whereinsaid first pass is positioned upstream of said second pass in saidbypass duct.
 30. The intercooling system as set forth in claim 29wherein said at least one of said more upstream locations is in a highpressure compressor.
 31. The intercooling system as set forth in claim21, wherein said at least one of said more upstream locations is in ahigh pressure compressor.
 32. The intercooling system as set forth inclaim 21, wherein said at least one of said more upstream locations isin a low pressure compressor.
 33. The intercooling system as set forthin claim 21, wherein a main fan delivers bypass air into a bypass ductand into said main compressor section and said heat exchanger positionedwithin said bypass duct to be cooled by bypass air.
 34. The intercoolingsystem as set forth in claim 21, wherein said cooling compressorincludes a centrifugal compressor impeller.
 35. A gas turbine enginecomprising: a heat exchanger for cooling air drawn from a portion of amain compressor section at a first temperature and pressure for coolingthe air to a second temperature cooler than the first temperature; acooling compressor compressing air communicated from the heat exchangerto a second pressure greater than the first pressure and communicatingthe cooling air to a portion of) at least a turbine section; said heatexchanger having at least two passes, with a first of said passespassing air in a direction having at least a radially outward component,and a second of said passes returning the air in a direction having atleast a radially inward component to the compressor; said first pass ispositioned downstream of said second pass in said bypass duct; a mainfan delivers bypass air into a bypass duct and into said main compressorsection and said heat exchanger positioned within said bypass duct to becooled by bypass air; said cooling compressor includes a centrifugalcompressor impeller; and said at least one of said more upstreamlocations is in a low pressure compressor.
 36. The gas turbine engine asset forth in claim 35, wherein an auxiliary fan is positioned upstreamof the heat exchanger.
 37. The gas turbine engine as set forth in claim35, wherein said main fan is driven by a turbine rotor through a gearreduction.
 38. The gas turbine engine as set forth in claim 37, whereina bypass ratio is greater than
 10. 39. The gas turbine engine as setforth in claim 38, wherein a gear ratio of said gear reduction isgreater than 2.3.
 40. The gas turbine engine as set forth in claim 37,wherein a gear ratio of said gear reduction is greater than 2.3.